This invention relates generally to gas turbine engines and more particularly to shroud assemblies utilized in the high pressure turbine section of such engines.
It is desirable to operate a gas turbine engine at high temperatures most efficient for generating and extracting energy from these gases. Certain components of a gas turbine engine, for example stationary shroud segments which closely surround the turbine rotor and define the outer boundary for the hot combustion gases flowing through the turbine, are exposed to the heated stream of combustion gases. The base materials of the shroud segment can not withstand primary gas flow temperatures and must be protected therefrom
Impingement cooling on the back side and film cooling on the hot flow path surface are the typical prior art practices for protecting high pressure turbine shrouds. The film cooling effectiveness on the shroud gas path surface is typically not high because the film is easily destroyed by the passing turbine blade tip. Another method to keep the shroud temperature low is to apply a layer of thermal barrier coating (“TBC”) on the hot flow path surface to form a thermal insulation layer. One particular effective kind of TBC is dense vertically microcracked TBC or “DVM-TBC”. To prevent spalling of the TBC, the temperature of the underlying bond coat must be kept below about 950° C. (1750° F.). Furthermore, drilling cooling holes through a TBC can damage the structure of the TBC and result in spallation. Certain prior art shrouds with a DVM-TBC have a sufficient operational life without film cooling. However, engines are now being designed to be operated at high temperatures for extended periods of time, requiring both a TBC coating and effective cooling.
Accordingly, there is a need for a turbine shroud which can provide film cooling coverage over the flow path surface without causing spallation of a coating applied thereto.